This invention relates to axial flow gas turbine engine compressors, and specifically to a means for enhancing stall margin of the compressor without adversely impacting efficiency by including a treatment to the case of the compressor.
In an axial flow gas turbine engine 10, such as the type used on aircraft and shown in FIG. 1, air is compressed in a compressor section 12, mixed with fuel combusted in a combustor section 16, and expanded through a turbine section 14 that, via one or more shafts, drives the compressor section 12. The overall efficiency of such engines is a function of, among other factors, the efficiency with which the compressor section 12 compresses the air. The compressor section 12 typically includes a low pressure compressor 18 driven by a shaft 26 connected to a low pressure turbine 24 in the turbine section 14, and a high pressure compressor 20 driven by a shaft 26 connected to a high pressure turbine 22 in the turbine section 14. The high and low compressors 18, 20 are multi-stage where the air flows in the axial direction through a series of rotating blades and stationary stators or vanes that are concentric with the axis of rotation (longitudinal axis). Each stage includes a row of blades and a row of stators.
The stages are arranged in series, and as air is compressed through each stage, the air experiences an incremental increase in pressure. The total pressure increase through the compressor is the sum of the incremental pressure increases through each stage, adjusted for any flow losses. Thus, in order to maximize the efficiency of the gas turbine engine 10, it would be desirable, at a given fuel flow, to maximize the pressure rise (hereinafter referred to as xe2x80x9cpressure ratioxe2x80x9d) across each stage of the compressor.
One of the design considerations facing designers of axial flow gas turbine engines is a condition known as compressor stall. Compressor stall is a condition in which the flow of air through a portion of a compressor stage ceases, because the energy imparted to the air by the blades of the compressor stage is insufficient to overcome the pressure ratio across the compressor stage. If no corrective action is taken, the compressor stall may propagate through the compressor stage, starving the combustor section of sufficient air to maintain engine speed. Under some circumstances, the flow of air through the compressor may actually reverse direction, in what is known as a compressor surge.
Compressor stalls in the high compressor are of great concern to engine designers, and while compressor stalls can initiate at several locations within a given stage of a compressor, it is common for compressor stalls to propagate from the blade tips where vortices occur. It is believed that the axial momentum of the airflow at the blade tips tends to be lower than at other locations along the airfoil. From the foregoing discussion it should be apparent that such lower momentum could be expected to trigger a compressor stall.
As an aircraft gas turbine engine accumulates operating hours, the blade tips tend to wear away the tip shroud, increasing the clearance between the blade tips and the tip shroud. As those skilled in the art will readily appreciate, as the clearance between the blade tip and the tip shroud increases, the vortices become greater, resulting in a larger percentage of the airflow having the lower axial momentum discussed above. Accordingly, engine designers have sought to remedy the problem of reduced axial momentum at the blade tips of high compressors.
Treatment of the casing to enhance the stall margin and, more specifically, to desensitize the high pressure compressor 20 of an engine 10 to excessive clearances between the blade tips and tip shrouds (tip seal or outer air seal) is shown and described in commonly assigned U.S. Pat. Nos. 5,282,718 to Koff et al., 5,308,225 to Koff et al., 5,474,417 to Privett et al., 5,607,285 to Bryne et al. In practice, the above referenced patents include a plurality of baffles or vanes in the shroud assemblies of the blades. Although effective, these vanes and baffles require an additional machining operation to manufacture. Also, the prior designs require that the passage is formed by mechanically joining the outer diameter and the inner diameter. Thus, manufacturing time and cost is increased to produce the prior art casing treatments. Further, the prior art casing treatments are contained entirely in the blade tip shroud thus requiring yet another machining operation through any abradable material that is used for sealing.
Thus, what is needed is a casing treatment which prevents compressor stall, eases part producibility and reduces manufacturing costs, while increasing the maintainability, assembly and safety as compared to the prior art.
The above discussed and other drawbacks and deficiencies are overcome or alleviated by the present invention.
The assembly of the present invention is formed by a single integral casting that enhance, compressor stall margin, and reduces manufacturing costs while increasing the maintainability, assembly and safety, as compared to the prior art.
In an exemplary embodiment of the present invention, a cast assembly for a gas turbine engine compressor comprises a stator outer platform, a boss, a passage, a blade outer shroud and a circumferential slot. A boss is formed in the platform and includes a passage. The passage includes an inlet and an outlet. The inlet is disposed downstream and proximate to the leading edge of the blades. The outlet being disposed upstream of the leading edge of the blades. The passage circumferentially converges inward from the inlet to the outlet. The outer shroud joins with said outer platform during the assembly of the compressor such that the circumferential slot is formed. The circumferential slot is in flow communication with the inlet.
The foregoing and other features and advantages of the present invention will become more apparent from the following description and accompanying drawings.